Combustion system for thermal powerplant



. 1, 1959 E. wou.

coususuou SYSTEM FOR THERMAL POWERPLANT 2 Sheets-Sheet 1 Original FiledMarch 10. 1949 Inventor: Edward Woll, y MW I FLOW UNIFORM FL 0 W REGIONrkhuslnou REGION FLOW His Attorney.

Dec. 1, 1959 E. woLL COMBUSTION SYSTEM FOR THEPZMAL POWERPLANT 2Sheets-Sheet 2 Original Filed March 10. 1949 His Attorney.

United States Patent COMBUSTION SYSTEM FOR THERMAL POWERPLANT EdwardWoll, Wenham, Mass., assignor to General Electric Company, a corporationof New York Continuation of application Serial No. 80,696, March 10,1949. This application October 24, 1955, Serial No. 542,391

3 Claims. (Cl. 60-35.6)

This application is a continuation of my prior application, Serial No.80,696 filed March 10, 1949, now abandoned. 1

This invention relates to internal combustion powerplants and hasparticular reference to the combustion of hot gases subsequent to theirdischarge from a turbine stage in such a plant. It has found particularutility in connection with gas turbine powerplants for effecting thepropulsion of aircraft, as for example in jet-propelled aircraft. It isthis application of the invention which I have elected specifically toillustrate and describe. It is to be understood, however, that theinvention is not limited thereto necessarily.

A gas turbine power plant for the propulsion of aircraft may include anair compressor, combustion apparatus, and a gas turbine compactlyarranged in series-flow relation to keep the weight, overall length anddiameter to a minimum. In such a power plant, the turbine is driven byhot gases generated by the compressor and the combustion apparatus, andthe turbine extracts at least suflicient power from these gases to drivethe compressor. The remaining power in the hot gases may be used topropel the aircraft by ejecting the gases rearwardly from the turbine athigh velocity through a suitable propelling nozzle.

In numerous instances it has been found that when powerplants of thetype described and installed in a jetpropelled airplane, the powerplantwill barely yield adequate thrust for take-off purposes, even though thepowerplant is capable of furnishing the thrust required for maintainingnormal flight speeds. Experience has also indicated that after amilitary airplane is put into actual field service, its take-off thrustmay have become inadequate due to increases in airplane weight which hasbeen brought about by various military requirements. In many instancesit has also been found that combat military aircraft are invariablycalled upon to do more than that for which they were originallydesigned. Special military missions often demand extra bursts of speedfor short periods of time. In view of these special instances, and forthe additional reason that jet-propelled aircraft have relatively poortake-off and climb characteristics resulting from inherently lowpropulsion efiiciency of such aircraft at low flight speeds, it isdesirable that additional thrust be made available without resorting tothe use of an oversize powerplant with the attendant increase in deadweight.

An elementary way to augment the normal thrust of such a powerplant issimply to operate at an over-speed condition. However, this has theserious disadvantage of appreciably reducing the operating life of thepowerplant. Another simple way in which thrust augmentation may beachieved is by the injection of water, alcohol, ammonia, or othersuitable fluids at the inlet of the compressor, or by injection of suchfluids into the combustion chambers, but experience has shown that eachof these relatively simple expedients is subject to the criticism thatonly limited increases in thrust can be obtained. The

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peak pressures that can be supported by the compressor impose a limit onthe rate of fluid injection, and exceeding this limit would cause thecompressor to stall, thus leading to operation of the turbine and thecombustion chamber parts at temperatures greatly exceeding their normaland safe operating values. In addition, experience has shown thatoperation at high rates of fluid injection has led to decreased life ofthe power plant. Many methods for achieving thrust augmentation areknown but are subject to the above-mentioned or similar limitations.

Accordingly, it is an object of this invention to provide a novelarrangement for augmenting the power developed by a gas turbinepowerplant which overcomes the abovementioned disadvantages.

A further object is to provide a system of the type described, which ismechanically simple yet thermodynamically effective, and which does notimpair or interfere with normal powerplant performance during periodswhen the augmentation system is not in use or operation.

Another object is to provide an after-burning or reheating arrangementfor a powerplant of the type described which is capable of operatingwith stable combustion over a wide range of fuel flows.

Still another object of the invention is to provide improved means forigniting combustible fluids which are flowing at velocities exceedingthe velocity of flame propagation.

It is a further object of the invention to provide new and improvedmeans for operating turbine apparatus whereby ignition and stablecombustion of fluids flowing at velocities exceeding the velocity offlame propagation can be eifected over a wide range of fuel flow rates.

Other objects and advantages will be apparent from the followingdescription taken in connection with the accompanying drawings, in whichFig. l is a half-sectional view of an exhaust system for a gas turbinepowerplant in accordance with the invention; Fig. 2 is a diagrammaticrepresentation of fluid flow in the region of a sudden enlargement in aflow conduit; Fig. 3 is an enlarged detailed view illustrating thenature of the flow pattern of the hot gases in the region of a suddenenlargement provided in the exhaust system of Fig. 1; Fig. 4 is anenlarged detailed view showing a modified nozzle arrangement forinjecting fuel into the fluid passage; and Fig. 5 is a modifiedembodiment of the arrangement shown in Fig. 1.

Referring now to Fig. 1, a turbine inlet casing is partially indicatedat 1, having an inlet passage 2 which communicates with turbine nozzles3. Motivevfluid under pressure and at high temperature is supplied tothe turbine nozzles 3 from inlet passage 2. For example, the motivefluid may be air which is compressed in a suitable compressor and isheated subsequent to its discharge from the compressor by burning fuelin the air in suitable combustion chambers. When air is employed as amotive fluid, it is common practice to employ high air-fuel ratios, forexample, of the order of 60 to l in order to limit the maximumtemperature of combustion to a value which the turbine and combustionchamber materials can withstand with a reasonable degree of safety. Itwill be appreciated by those skilled in the art that in such cases themotive fluid discharged from the turbine will contain sufficientunburned oxygen to support additional combustion. The compressor,combustion system, and connecting conduits are not material to anunderstanding of the present invention, and are more particularlydisclosed in US. Patent 2,432,359Streid, and in US. Patent2,71l,074Howard, assigned to the assignee of the present application.The turbine nozzles 3 expand the fluid to a suitable velocity and directit at a proper angle to a I turbine wheel or bladed rotor 4 which isrotatably sup1 ported in suitable bearings (not shown) in the turbinecasing. The turbine rotor 4 is connected by means of suitable shaftingand/or gearing (not shown) to a consumer of power, for example, to thecompressor which furnishes the pressurized motive fluid.

After passing through blading 5 of the turbine wheel, the fluid isdischarged to an exhaust passage 6. It has been found convenient,particularly in aircraft service, to construct exhaust passage 6 by anouter wall or casing 7 and an inner Wall 8, each of which walls may becones or portions thereof and made of relatively thin sheet metal.Supporting and flow straightening members 9 may be provided for thepurpose of maintaining walls 7, 8 concentric and in spaced relation toeach other to define an annular passage, as well as for the purpose ofremoving tangential components of elocity from the fluid discharged fromthe turbine blades 5. it is to be understood, however, that the abovedescribed configuration is not limiting and that other a'rangements andwall shapes may be employed with equally good results. A flanged portionis provided at one extremity of wall 7 to facilitate attachment of theexhaust structure comprising walls 7, 8 and supporting members 9 toturbine casing 1, which is provided with a similar mating flange 10'.Flange .10, and thus the entire exhaust structure, is secured to themating flange 10' on the turbine casing by a quick-disconnect clampwhich may be of the type disclosed in U.S. Patent 2,424,436-Crater, orby other types of known securing means such as threaded fastenings.Similarly, flange 11 is provided at the opposite extremity of wall '7for attachment to a similar mating flange 11' secured to a conduit 12which conducts the gases to a suitable propelling nozzle 13.

Since, as previously indicated, the turbine removes at least sufiicientpower from the net motive fluid to drive the compressor, the pressureand temperature of the motive fluid discharged from the turbine wheel topassage 6 are considerably less than the pressure and temperature of thefluid in passage 2 preceding the turbine nozzles 3. in power plants usedfor effecting the propulsion of aircraft, the total or dynamic pressureof the motive fluid in passage 6 normally will be greater than thestatic ambient atmospheric pressure. Expansion of the motive fluid fromits super-ambient pressure and temperature in exhaust passage 6 to thestatic pressure of the ambient atmosphere through nozzle 13 produces acertain velocity of the motive fluid in a rearward direction relative tothe aircraft. The reaction resulting from this reaward velocity producesa forward thrust on the aircraft.

It is well known that the magnitude of thrust reaction in a dynamicfluid system is proportional to the mass fiow of fluid and also to thevelocity of the fluid. Thus, if either the mass flow or the velocity ofthe fluid can be increased in any manner, an increased thrust willresult. It is also well known that the velocity of fluid issuing from anozzle can be increased by increasing the pressure or temperature of thefluid preceding the nozzle. To this end I provide means for the additionof heat to the motive fluid at a location between the entrance to theannular exhaust passage 6 and the entrance to nozzle 13. A plurality ofspray nozzles 14 are provided at the downstream side of the turbinewheel 4 and are arranged to inject fuel into exhaust passage 6. Spraynozzles 14 may be of any type which produce a fine spray of fuelparticles 14', and may be similar to that disclosed in U.S. Patent2,524,820-Miles assigned to the same assignee as the presentapplication. In order to minimize the pressure loss in the exhaustpassage 6, it is desirable to arrange the spray nozzles 14 in such amanner that they do not extend into the stream of fluid. This may beaccomplished by providing recessed portions 15 in wall 8 of suflicientdepth to permit the installation of nozzles 14 in substantially flushrelation with the wall of passage 6. It is to be understood, however,that other arrangements such as that shown in Fig. 4 may be used withgood results.

From the standpoint of obtaining maximum mixing of the fuel and hotgases prior to ignition, it is desirable to arrange nozzles 14 in acommon plane axially spaced downstream from the turbine wheel and withsuch axial spacing reduced to an absolute minimum. Due to manufacturingvariations, it has been found to be virtually impossible to achieveabsolute uniformity of pressure distribution in passage 6 immediatelyadjacent the turbine wheel and thus there may be a crosscirculation ofhot gases across the rear face of turbine wheel 4. It is thereforedesirable to locate nozzles 14 in a common plane which is more remotelyspaced from the turbine wheel in order to preclude the possibility offuel particles, which may ignite spontaneously, becoming entrained inthe above-mentioned cross-circulation. In practice, the design is made acompromise of these conflicting considerations, and experiments haveindicated that the exact location of the plane, in which nozzles 14 arelocated, does not appear to be critical.

in order to achieve uniform heating of the fluid and thus to avoidexcessive distortion and perhaps even failure of the exhaust structure,nozzles 14 are uniformly spaced around the outer periphery of inner wall8 and are connected to a common manifold 16 by connecting conduits 17.Fuel under pressure is supplied at a variable rate from a suitablesource (not shown) by means of conduit 13 to manifold 16 and thus to thespray nozzles 14. In aircraft service it is necessary to operate thepower plant at sea level and at high altitude, and under such varyingconditions of operation the air flow "ray vary through a range of atleast 10-20 to 1. Therefore, it is necessary to provide for a range offuel flow of the same magnitude. In addition, a reheat system foraircraft service must be capable of stable operation for fuel flows bothequal to and considerably less than the theoretical stoichiometricproportions of fuel and air for complete combustion of the fuel. In viewof these considerations and to keep the fuel pressure within practicallimits, it may be desirable to employ nozzles 14 of the duplex type andprovide two manifolds 16 which are connected by a flow divider (notshown). A duplex nozzle arrangement of the type suggested is more fullydescribed in U.S. Patent 2,622,393Edwards et al. assigned to theassignee of the present application. As indicated in Fig. 1, thesupporting members 9 may be hollow. Such arrangement furnishesconvenient means for introducing fuel to the interior of wall structure8 and to the spray nozzles 14, and at the same time the hollowconstruction of member 9 effects considerable saving in weight, which isparticularly important in aircraft service.

For reasons which will appear later, it is necessary to stall the fluidor a portion of the fluid flowing from the exhaust passage 6 to conduit12. The term stall as used herein means that the fluid velocity isreduced to a low value less than the velocity of flame propagation inthe fluid. The term stalled region or quiescent zone as usedhereinafter, refers to a region within the fluid passage wherein thefluid is stalled as previously defined. Extensive testing experience hasshown that when fuel is introduced into a high velocity fluid stream,even though in the proper stoichiometric proportions to form acombustible mixture, the mixture may be very difficult to ignite.Experimentation has further shown that in order to effect ignition andto obtain burning under stabilized conditions, it is necessary to stalla portion of such a high velocity fluid stream, or at least to reducethe velocity of av portion of the combustible mixture to a very lowvalue as compared to the velocity of the fluid in other portions of thepassage. One very simple way in which such a stalled region may becreated is to insert an obstruction in the fluid passage. in such casethe region immediately adjacent the obstruction at its downstream sideWill be effectively stalled and the local velocities of the fluid Withinthe region will be extremely low compared to the velocity of the fluidin other portions of the passage. Obviously such a method for effectingignition in a thermodynamic fluid system is subject to the seriousobjection that it introduces pressure losses into the system therebyadversely elfecting thermodynamic efliciency and greatly reducing themechanical power which might otherwise be obtained from such a system.

I have found that such a stalled region can be created by providing awall 8', which effectively cuts oil the vertex of the cone which wouldotherwise be formed by wall 8. As illustrated in Fig. 1, wall 8'preferably is hemispherical in shape, although it may be constructed ofother suitable shapes. Considering now, the geometry of cone portiondefined by wall 7, 8 it will be seen that the effect of cutting ofi? thevertex of the cone, which would otherwise be formed by wall 8, is tocreate a sudden enlargement in the area of the fluid passage 6 t0 theright of plane 19. In other words, to the left of the plane 19, exhaustpassage 6 is the annular space defined by walls 7, 8, while to the rightof plane 19 the above-mentioned passage is circular in cross section andis defined first by wall 7 alone, and then by Wall 12 alone.

It will be seen presently that this change in area produces twoimportant effects. First, the increased area of the passage will reducethe velocity of the fluid in the stalled region to a certain velocityless than its original value. The magnitude of the reduced velocity willbe determined substantially by the cross-sectional area of conduit 12and the average density of the fluid therein. Secondly, in the immediatevicinity of the sudden enlargement, or more specifically, in the centralportion of the passage near the righthand extremity of wall 8,

the fluid flowing in the passage will undergo a transition in which itspressure and velocity rapidly change in an attempt by the fluid tocompletely fill the cross-sectional area of conduit 12. In thistransitional region, because of the sudden enlargement in flow area andthe attempt of the fluid to completely fill the available space, therewill be considerable turbulence and eddying of the fluid.

However, as will appear presently, the velocity of the fluid in thisregion will be relatively low compared to the velocity of the fluid inother portions of the system for example, in passages 6 or in conduit12.

A better understanding of the nature of the flow pattern established inthe region of the previously described sudden enlargement can be had bycomparison with similar flow patterns in the region of suddenenlargements as they are conventionally represented. Fig. 2 representsthe type of fiow pattern which is obtained when fluid which is flowingfrom left to right in a first conduit 20 flows into a second conduit 21having a substantially greater flow area than that of the first conduit.In an ideal case where there are no losses in the system, the totalpressure of the fluid would be the same on either side of plane 19' inwhich the sudden enlargement takes place. To the left of plane 19 thestatic pressure will be low relative to the static pressure at theright, and the velocity to the left of plane 19' will be considerablyhigher than that at the right due to the difference in flow area. Inthetransitional region indicated in the drawing, the velocity of the fluidis reduced and the static pressure is increased in the attempt of thefluid to completely fill the enlarged flow area. Due to the increasedstatic pressure at the right of plane 19 and because there is a tendencyfor a portion of the flow to separate at its periphery, a reverse flowor back eddy as indicated by arrows 22 will be formed. The regionsindicated by arrows 22, constitute the stalled regions of the typepreviously defined. The portion of the total flow, which separates andflows in the manner indicated by arrows 22, usually is a smallpercentage of the total flow and the velocities in the stalled regionsare relatively low due to the fact that the differences in staticpressure at locations to the left of plane 19' and at the right of plane19' generally are small.

Referring now to Fig. 3, the main flow from passage 6 to conduit 12 isindicated by arrows 23. As previously indicated, the effect of cuttingofi the vertex of the cone which would otherwise be formed by wall 8 isto create a sudden enlargement in the passage at plane 19. Thetransitional region in which the flow undergoes a change in an attemptto completely fill the increased area of the passage is indicatedbetween planes 19 and 19". For the same reasons as already indicated inconnection with Fig. 2 the static pressure to the right of plane 19 isgreater than the static pressure of the fluid at plane 19. The higherstatic pressure at the right of plane 19 causes a portion of the flow toseparate and form. a reverse eddy indicated by arrows 24. Thus it willbe seen that the stalled region is bounded on the left by wall 8', andto the right of plane 19 substantially by the vertex of the cone surfacewhich would be formed by the continuation of wall 8 as indicated bybroken lines 8". Experiments have shown that the ratio of the projectedarea defined by wall 8' to the flow area of conduit 12 including thestalled region should be maintained between .05 and .15 to obtainsatisfactory and stable combustion at average fluid velocities of theorder of 500 feet per second or higher. Stated in another way, the flowarea of conduit 12 should be from 5 to 20 percent greater than the areaof annular passage 6 at plane 19.

In gas turbines of the type described, the velocity of the fluid leavingthe turbine wheel may be of the order of 800 to 1200 feet per second.Velocities ofthis magnitude generally exceed the velocity of flamepropagation in the fluid. It has been shown by mathematical analysis andverified by experiments that the pressure losses in a high velocityfluid system which arise solely from aerodynamic and thermodynamicreasons may greatly exceed pressure losses which are due to fluidfriction and to fluid shearing forces. This is particularly true whereextremely high velocities are involved. It can also be shown bymathematical analysis that there is a decrease in total pressure uponthe application of heat to a stream of flowing fluid when the heat isapplied at constant pressure or at constant flow area, and that thedecrease in total pressure can be minimized by reducing the Mach numberof the fluid prior to the application of heat to as low a value as ispracticable. It will be appreciated by those familiar with the art thatthe Mach number is the ratio of the actual local velocity of the fluidto the local acoustic velocity in the fluid. Therefore, it is common toprovide an exhaust passage 6 Which gradually increases in area in orderto diffuse and thus reduce the velocity of the fluid to as low a Machnumber as practicable before introducing the gas to the stalled region.It will be obvious to those familiar with the art that this requirementcalls for extremely large passage areas and, particularly in aircraftservice, a compromise must be made in order to minimize pressure dropand-at the same time obtain apparatus which is reasonable in size andweight. Experiments have shown that good results are obtained withreasonable passage areas if the average fluid velocity is reduced to aMach number between .2 and .3 at the upstream side of plane 19. It willalso be obvious that by providing exhaust passage 6 with a graduallyincreasing area, a gentle transition from the relatively small areadefined by the turbine annulus to the relatively large area defined byconduit 12 is provided which further tends to minimize pressure loss.The function and the effect upon the operation of the system illustratedin Fig. 1 of the stalled region referred to above will appear later. Itwill sufiice at this point to state that by means of this 7 stalledregion, ignition of the fuel in the high velocity stream of fluidsubsequent to its discharge from the turbine blades is accomplished.

It is desired particularly to point out that suflicient structure hasbeen disclosed at this point to effect ignition of a stream of flowingfluid which has been mixed with fuel in suflicient proportions to createa combustible mixture. It is necessary, however, that the averagetemperature of the stream of flowing fluid be at least as high as theignition temperature of the combustible mixture. In cases where thiscondition cannot be met, I have found that combustion may be initiatedin the stalled region by alternative methods to be described later.

Combustion of the high velocity fluid-fuel mixture is initiated in thefollowing manner. Motive fluid is supplied to the turbine inlet passage2 under pressure and at high temperature. The fluid is expanded by meansof nozzles 3 which direct the fluid at suitable velocity and in theproper direction to the turbine blading 5. The turbine is driven by thehot motive fluid and at least sufficient power is extracted therefrom todrive the compressor. Upon being discharged from the turbine wheel, themotive fluid is discharged into passage 6 at reduced pressure andtemperature, and at high velocity. As previously indicated, it is commonfor the axial velocity of the fluid in passage 6 to be of the order of800 to 1200 feet per second. As previously indicated, fuel is introducedinto the high temperature fluid flowing in passage 6. The fuel isintroduced at low velocity relative to the velocity of the fluid streamand in a substantially radial direction as substantially normal to thepath of the high velocity fluid stream, to insure adequate mixing of thefuel and the hot motive fluid. Because this fuel is introduced at lowvelocity there is litttle tendency for it to penetrate deeply into thefluid stream. A stratified layer of mixed fuel and hot motive fluid isformed adjacent to inner wall 8. The combination of the relatively lowradial velocity of the fuel and the high axial velocity of the hotmotive fluid in passage 6 tends to confine the fluid-fuel mixture to theregions immediately adjacent wall 8, and also tends to confine thismixture to the central portion of conduit 12. When the fluid-fuelmixture passes to the right of plane 19, at least a portion of themixture separates and enters the stalled region in the manner previouslydescribed. As previously indicated, the velocity of the mixture in thisregion will be greatly reduced from its previous value, and when thetemperature of the fluid is above the ignition temperature of thefluid-fuel mixture spontaneous ignition will occur if the localvelocities of the mixture are less than the velocity of flamepropagation. After combustion has beeninitiated, burning is firstconfined to the stalled region previously defined in connection withFigures 2 and 3, and then tends to propagate in all directions. Thatportion which has been ignited inthe stalled region serves as a pilotburner to ignite the previouslydescribed stratified mixture whichcontinues to flow along the Wall 8 and into the central portion ofconduit 12. Ignition of this mixture first occurs at the point marked Ain Fig. 1. Since the speed of flame propagation is less than the axialvelocity of the fluid-fuel mixture, the flame front will be deflectedfrom a radial plane to form a combustion zone at the center of conduit12 substantially as indicated by broken lines 25. In certain types ofservice the temperatures within the combustion zone defined by litres 25may be as high as 3800 F., and thus the importance of keeping the flameconfined to the central portion of the pipe will be appreciated. It isan important feature that according to the invention the flame isconfined to the central portion of the exhaust conduit 12 by the highvelocity gas stream, and that a relatively cool annular layer of motivefluid is interposed between the flame boundary and the inner wall ofconduit 12, thereby maintaining the conduit walls at a temperature muchlower than that in the central portion of the gas stream. Experimentshave shown that good results are obtained by limiting the penetration ofthe fuel particles to a value not exceeding percent of the spacingbetween the inner and outer Walls 7 and 8.

Again referring to Fig. 1, if desired, an additional fuel spray nozzle26 may be provided at the central axis of wall 8. In such case nozzle 26serves as a pilot burner and is supplied with fuel conveyed by a suitaleconduit 27 from the fuel supply means. The fuel conveniently may betaken from manifold 16 as indicated or may be supplied from an entirelyseparate source (not shown).

Still referring to Fig. l, in applications where the temperature of thegases discharged from the turbine blade 5 is less than the ignitiontemperature of the fluidfuel mixture, other means must be resort-ed toin order to initiate combustion of the mixture. T 0 this end an ignitingdevice 28, which conveniently may be a suitable electrically energizedspark plug, is provided to initiate combustion in the stalled region.Spark plug 28 is supported on outer wall 7 by a mounting flange 29 insuch a manner as to project through walls 8 and 8'. The spark plug islocated in relation with wall 8' so that electrodes 39 defining thespark gap are located within the space enclosed by wall 8' and plane 19.Electrodes 30 may be arranged so that the spark gap is in the normalpath of the fuel particles discharged from nozzle 26 althoughexperiments have shown that the location of the spark gap is notcritical. A cylindrical coaxially spaced metal sleeve 31 surrounds thespark plug in order to protect it from the harmful effect of the hotgases flowing inpassage 6.

In operation, the arrangement shown in Fig. 1 functions in a manner verysimilar to the arrangement previously described wherein nozzle 26 andspark plug 28 are omitted. Fuel is introduced into passage 6 by nozzles14- in such a manner as to limit the degree of penetration of the fuelparticles into the gas stream. As previously indicated, this causes astratified layer of fluidfuel mixture to be formed adjacent to wall 8.As before, a portion of this mixture becomes stalled in the stalledregion with the local velocities of the mixture in this region less thanthe velocity of flame propagation so that the combustible mixture can beignited by energizing the spark plug. It will be apparent that the sparkplug may be energized either manually, or automatically when fuel issupplied to nozzles 14. Once ignition has been initiated in the stalledregion, combustion occurring Within this region serves as a pilot burnerto initiate and stabilize combustion of the stratified layer of fuel gasmixture which flows past point A in the manner previously described.

More accurate control of the degree of penetration of the fuel spray 14'into exhaust passage 6 may be had at the expense of a slight amount ofpressure loss by employing a spray nozzle 14 of the type indicated inFig. 4. The body of the nozzle is made of suificient length to securethe desired degree of penetration of fuel particles and a plurality oforifice openings 32 in a common plane and normal to the axis of thenozzle body are provided therein. It will be apparent that the spacingof openings 32 with respect to inner wall 8 will afford accurate controlof the degree of penetration of the fuel particles into the passage.

In operation, the openings 32 are arranged to discharge into passage 6in a direction normal to the axis of the passage and tangentially Withrespect to inner wall 8. Since the fluid flow in passage 6 ispredominantly axial in direction, good mixing of the fuel and motivefluid is obtained and a stratified layer of fluid-fuel mixture is formedadjacent to wall 8 in the manner previously described.

A modified embodiment of the invention is illustrated in Fig. 5 whereinan arrangement is disclosed for effecting ignition of the fuel-gasmixture when the temperature of the gas flowing in passage 6 is lessthan the ignition temperature of the fuel-gas mixture and when it isdesired to eliminate the need for an electric igniting device. A conduit33 connects the passage 2 preceding the turbine nozzles and the stalledregion. A flow restricting orifice 34 is provided in conduit 33 to limitthe flow of fluid through the conduit and to reduce the static pressureof the fluid within conduit 33 and of the downstream side of orifice 34to a value only slightly in excess of that existing in the stalledregion. The reduction of pressure of the fluid flowing through conduit33 is necessary to insure that this fluid is introduced into the stalledregion at relatively low velocity. Since no power is removed from thegases flowing through conduit 33 in this manner, it will be obvious thatthe temperature of this fluid will greatly exceed the temperature of thefluid discharged by the turbine into passage 6.

In operation, motive fluid is supplied to the turbine inlet passage 2under pressure and at high temperature. A portion of the fluid isexpanded by means of turbine nozzles 3 which direct the fluidat-suitable velocity and in the proper direction relative to the turbineblading 5 to drive the turbine. Other portions of the motive fluid flowthrough conduit 33 and are introduced into the stalled region atrelatively low velocity. The temperature of motive fluid thus introducedto the stalled region will greatly exceed the temperature of the fluiddischarged by the turbine for the reason previously indicated. Fuel isinjected into exhaust passage 6 in the manner previously indicated inconnection with Fig. 1 and 4 so as to form a stratified layer offluid-fuel mixture surrounding wall 8. In general, the temperature ofthe fluid from conduit 33 will exceed the ignition temperature of thefluid-fuel mixture. As in the preceding cases, a portion of thisstratified layer enters the stalled region and the remaining portionflows into the central portion of conduit 12. That portion of themixture which enters the stalled region upon coming in contact with thehigh temperature gas introduced to the region by means of conduit 33 iscaused to ignite spoutaneously due to the low velocity of the mixture inthis region and the high temperature of the gases introduced throughconduit 33. Combustion of the mixture in the stalled region in thismanner serves as a pilot burner to initiate and stabilize combustion ofthe remaining portion of the stratified layer of gas-fuel mixtureflowing past point A in the drawing. The shape of the combustion zoneand the confining of flames to the central portion of conduit 12 as wellas cooling of the walls of conduit 12 are eflected in the same manner asin the previous arrangements.

Thus it will be apparent that the invention provides a novel andrelatively simple system for augmenting the power developed by gasturbines wherein the flame is confined to the central portion of thedischarge conduit and a relatively cool layer of motive fluid isinterposed between the flame boundary and the conduit wall to maintainthe wall at a temperature much lower than that in the central portion ofthegas stream. Furthermore the invention permits stable combustion overa wide range of fuel-flows; the use of high combustion temperatures inthe re-heat system is permitted because of the improved cooling; the useof light gauge sheet metal walls is permitted because the surface of theexhaust conduit is swept by turbine exhaust gases at substantiallyturbine discharge temperature; and the invention permits the use of asubstantially unobstructed exhaust passage, thereby minimizing systemlosses and reducing the possibility of mechanical failure; andperformance in unimpaired by the presence of such an arrangement duringperiods when the augmentation device is hot in operation.

While particular embodiments of the invention have 10 been illustratedand described, it will be obvious to those familiar with the art thatvarious changes and modifications may be made without departing from theinvention, and it is intended to cover in the appended claims all suchchanges and modifications as come within the true spirit and scope ofthe invention.

What is claimed is:

l. A reheat system for use in turbo-machines comprising: a passage fordilfusing a stream of combustion-supporting fluid and reducing itsvelocity to a Mach number of .25, said passage being defined bydiverging inner and outer walls, the inner wall forming a frustum of acone; fuel injection means mounted flush with the inner wall near theupstream end of the frustum, said injection means being adapted toinject fuel into the stream of fluid so as to form a stratum offuel-fluid mixture adjacent the inner wall surrounded by a stratum offuelfree fluid adjacent the outer wall; a re-entrant wall extending intothe downstream end of the frustum with its edge coincident with thedownstream end of the inner wall, the depth of the re-entrant wall beingno greater than the radius of the downstream end of the frustum, saidre-entrant wall cooperating with the outer wall to form a suddenenlargement in said passage for stalling a portion of the fuel-fluidmixture in the center of the enlargement; and ignition means mounted inthe re-entrant wall for igniting the stalled portion of the fuelfluidmixture, which in turn ignites the remainder of the mixture; whereby theburning fluid-fuel mixture is confined and surrounded by the stratum offuel-free fluid, thus preventing burning adjacent the outer wall of thepassage.

2. In a turbine power plant, reheat means for combustion-supportingfluid comprising: coaxial spaced inner and outer walls defining a fluidpassage having a diifusing portion of annular cross-section and adownstream portion of substantially circular cross-section with a flowarea at least 10% greater than that of said diffusing portion, saidinner Wall defining the frustum of a cone terminating abruptly at thechange from annular to circular cross-section; a re-entrant wallextending into the downstream end of the frustum with its edgecoincident with the downstream end of the inner wall, said re-entrantwall forming a stalled region beyond the end of said frusturn; fuelinjection means mounted in the inner wall near the upstream end of thefrustum for forming a stratified layer of preselected depth offluid-fuel mixture within the diffusing portion of the passage adjacentsaid inner wall and surrounded by an outer layer of combustible fluidsubstantially free from fuel; and igniting means mounted in there-entrant wall, whereby the fluid-fuel mixture in the stalled region iscaused to ignite, the combustion of the ignited stalled portion causingthe remainder of said stratified mixture flowing past the periphery ofsaid stalled region to become ignited.

3. Apparatus in accordance with claim 6 wherein said igniting meansincludes a conduit for introducing a combustible fluid into the centralportion of said stalled region at a velocity substantially less than thevelocity of flame propagation in the fluid-fuel mixture and at atemperature substantially greater than the ignition temperature of saidstratified fluid-fuel mixture, and an orifice structure mounted in saidconduit for reducing the velocity of said combustible fluid.

References Cited in the file of this patent UNITED STATES PATENTS2,404,335 ,Whittle July 16, 1946 2,409,176 Allen Oct. 15, 1946 2,479,777Price Aug. 23, 1949 2,575,682 Price Nov. 20, 1951 2,619,795 Drake Dec.2, 1952 2 ,639,581 Cohen et a1. May 26, 1953

